Integral booster/ramjet drive

ABSTRACT

A missile drive is provided comprising a single combustion chamber shared by a first, acceleration stage and a second, ramjet cruising stage, said chamber housing the solid propellant to be consumed during acceleration. At least one air inlet is opened at the end of the acceleration stage to allow an air flow to be introduced into the combustion chamber. Moreover, at least one additional exhaust outlet forming an additional nozzle is provided in the back of the combustion chamber and means are provided to seal off said additional exhaust outlet or outlets throughout the initial acceleration stage, such that said additional exhaust outlet or outlets contribute, together with the converging-diverging nozzle, to ejecting the exhaust gas during the ramjet cruising stage.

This invention relates to a ramjet missile propulsion system or drivewith a built-in acceleration engine or booster, said drive comprising asingle combustion chamber, shared by a first, acceleration stage and asecond, ramjet cruising stage, wherein is stored the solid propellantused for missile acceleration, and further comprising aconvergent-divergent nozzle optimally dimensioned for acceleration stagepropulsion and at least one air inlet designed to open after theacceleration stage is enable enough air to enter the combustion chamberto at least compensate the drag force on the missile by the high-speedejection of the gases resulting from the combustion of ram air with themissile's fuel payload.

The invention thus concerns the propulsion of tactical-type missileshaving a flight envelope within the atmosphere such that air can be usedas one of the combustion agents in a ramjet engine.

Specifically, the invention concerns rocket ramjets operating with twosuccessive propulsion modes, ie. a rocket engine acceleration stageusing a fuel and an oxidizer both of which are carried aboard themissile, and a ramjet-based cruising stage using air scooped from theatmosphere by the moving missile with only the fuel carried aboard themissile.

Problems arise in the practical construction of such rocket ramjets dueto the fact that the optimum nozzle for acceleration or boost must havea much narrower throat than the throat cross section of the nozzle mostsuitable for the subsequent ramjet cruising stage.

Solutions involving integrated ramjet and booster stages have beenprovided, using a single combustion chamber the inside wall whereof iscommon to both propulsion stages, said designs comprising a nozzleoptimized for the acceleration stage, which is ejected during thetransition phase between acceleration and cruising, to produce a largeropening in the tail end of the combustion chamber being better suited toramjet cruising.

Unfortunately, for certain flights such as for example short, or mediumrange flights, ejection of the booster nozzle is unacceptable, since itconstitutes a danger either for the missile operator or for friendlytroops and populations.

In order to avoid the latter inconvenience, some prior art designs havedone away with the booster nozzle. In such nozzle-less power plants,charging of the solid propellant required throughout the accelerationstage is arranged by providing a center channel and a divergent rearcone which together ensure stable combustion and axial thrust throughoutthe acceleration stage. Nevertheless, in this type of power plantcombustion pressure considerably decreases as the missile acceleratessuch that the mean specific impulse during the acceleration stage issubstantially downgraded compared with that which the same propellantwould have provided had the exhaust gas been ejected through a nozzle.

Considering the fact that the length of the combustion chamber isgenerally dimensioned by the space required for the booster propellantand that any increase in the mass of this solid propellant entails anequivalent decrease in the mass of the ramjet fuel, any downgrading ofperformance in the acceleration stage entails a reduction of cruise timeand consequently a proportional shortening of the missile's range.

It is the object of the present invention to remedy the above-mentioneddisadvantages by providing for the escape of combustion gases producedduring the acceleration stage through an actual nozzle having beenoptimized for this stage and by avoiding ejection of said nozzle whilstensuring its suitability for subsequent operation in the ramjet cruisingstage.

Accordingly, the invention provides a propulsion system or drive asdefined at the beginning of the foregoing, having, besides the specialair inlet or inlets, at least one additional exhaust outlet, in the formof an additional nozzle located in the downstream end of the combustionchamber, and means of closing said additional outlet or outletsthroughout the initial acceleration phase, such that said additionaloutlet or outlets can also contribute, together with the main nozzle, toexhausting or combustion gases during the ramjet cruising phase.

In accordance with the invention, the main nozzle designed for theacceleration stage is kept in place during the cruising stage, butduring said latter stage, only part of the exhaust gas is ejectedthrough said main nozzle and the remainder is ejected through one ormore additional outlets provided at the back of the combustion chamber,said outlets, like the previously-mentioned chamber air inlets, beingopened only once all of the solid propellant used for acceleration hasbeen consumed.

In a first embodiment of the invention, said additional outlet oroutlets are provided in the back wall of the combustion chamber in orderto provide a substantially axial additional exhaust.

In another embodiment, said additional outlet or outlets are provided inan annular, rear lateral part of the combustion chamber.

The propulsion system according to the invention is advantageouslyfurther provided with an external structure surrounding the combustionchamber substantially to the rear end of said chamber and extendingforward of the chamber at least partly along the length of the missilebody so as to provide a permanent air scoop at the front of the missileand ensure ram air flow to the combustion chamber via said air inletswhilst maintaining the casing of the combustion chamber under balancedpressure during the ramjet cruising stage, such that during saidcruising stage, only a part of the air taken in by the missile's airscoops is introduced into the combustion chamber via the air inlet holeswhile the remaining air flow is discharged behind the missile viapurposely provided outlets without having passed through the combustionchamber but having nevertheless been mixed with at least that portion ofthe exhaust gas exiting via the additional exhaust outlets.

Said external structure can be extended backwards beyond the main nozzleto enable mixing of the secondary air flow with all of the exhaust gasfrom the combustion chamber.

Splitting of the air flow into a portion αQa which is admitted by theair inlets and contributes to combustion and a portion (1-α)Qa whichconstitutes the cooling flow is determined as a function of the thrustrequired and of the temperature withstandable by the combustion chamberwall.

The combustion chamber is centered with the help of struts insertedbetween the casing of said chamber and said previously-mentionedexternal structure. In the embodiment with additional exhaust outletsprovided in the shell said struts can be shaped such that the crosssection of the air passage between the external structure and thecombustion chamber casing provides for sonic flow of both the airupstream from the additional exhaust outlets and the air and combustiongas mixture downstream from the additional exhaust outlets.

The combustion chamber is preferrably given a thin, heat-conductingwall.

A number of other special features may also be advantageously providedas follows.

Conical, ie. two-way tapering ramps are inserted between the externalstructure and the fuel tank upstream from the combustion chamber airinlets to define an annular air scoop.

During the acceleration phase, the missile's air scoops are not sealedoff downstream, thus allowing the ram air to be discharged through therear to reduce drag.

A substantially conical fairing connects the rear neck of the combustionchamber to the outlet plane of the tapered out section of the main axialnozzle to form a central body establishing a diffuser in which thesecondary air flow/exhaust gas mixture can expand.

In one specific embodiment, the air inlets and the additional exhaustoutlets are sealed off during the acceleration stage by means of plugsspecially designed to be ejected at the end of said acceleration stage.

Alternatively, the combustion chamber can be designed so as to beautomatically pushed back axially by the gas pressure at the end of theacceleration stage over a pre-established distance to clear the chamberopenings while arranging, on the one hand, a space between the fuel tankand the chamber such as to form the air inlets and, on the other hand, aspace between the chamber and the main nozzle such as to form theadditional exhaust outlets. Depending on the applications envisaged, thecombustion chamber's sliding action thus serves either to open the airinlets, or the additional outlets, or both the inlets and outletssimultaneously.

More specifically, the sliding combustion chamber is initiallyimmobilized by means of a shearable tie and the geometricalconfiguration of said chamber is such that at the end of theacceleration stage, the direction of the axial resultant of the pressureforces on the chamber reverses and said resultant becomes greater thanthe resistance of the shearable tie.

Other features and advantages of the invention will become more readilyapparent in reading the following description of a few selectedembodiments, with reference to the appended drawings in which:

FIG. 1 is a cross-sectional diagram, taken axially, of a firstembodiment of the propulsion system according to the invention;

FIGS. 2 through 5 are cross-sectional views taken along II--II,III--III, IV--IV and V--V respectively of FIG. 1;

FIG. 6 is a schematic illustration of the principle of a dual air flowadapted to a propulsion system according to the invention;

FIG. 7 is a full cross-sectional view, taken axially, of a missileaccording to a second embodiment of the invention, showing, in a firsthalf-cross-sectional view, the working during the acceleration stage,and in a second half-cross-sectional view, the working during thecruising stage; and

FIG. 8 is an enlarged view of part of the missile shown in FIG. 7.

This invention can be readily applied to the various configurations ofram rockets with integral boosters. In fact, the air scoops required forthe ramjet stage of operation can be arranged in a number of differentways, including at the front, annularly, in the nose, as well asventrally or laterally. Similarly, the fuel required for the cruisingstage can be either liquid or solid. In the latter case it can be placedin an auxiliary self-pyrolyzing fuel generator or in contact with theinside wall of the combustion chamber.

Referring now to FIG. 1, the front part of the missile equipped with thedrive according to the invention is not shown. Only the fuel tank 8 andthe combustion chamber 2 are shown, both surrounded by an externaltubular structure 1 extending toward the back of the missile beyond themain exhaust nozzle 3. Said comustion chamber 2, attached to saidexternal structure 1, contains the propellant 4 required for theacceleration stage (see the lower half-cross-sectional view of FIG. 1).As long as fuel is stored therein and throughout the acceleration stage,the chamber's air inlets 5 remain sealed by plugs. In the ramjet stage,the fuel stored in the auxiliary tank 8 located upstream from thecombustion chamber 2 is injected into the combustion chamber 2 and thecombustion air scooped by the moving missile is introduced into thecombustion chamber 2 via the inlets 5 which have now been opened (seethe upper half-cross-sectional view of FIG. 1).

The embodiment illustrated in FIGS. 1 through 5 is associated withfrontal air scoop designs with said air scoops being located either inthe nose cone or, in annular form, downstream from the equipment whichmust be readily accessible, namely the charge and the instrumentationcompartment. In both cases secondary air is taken in upstream from thecombustion chamber 2 and allowed to circulate about the outside wall ofcombustion chamber 2 as a whole, thus providing chamber cooling andpreheating of said secondary air at the same time.

More specifically, the embodiment shown in FIG. 1 is provided with anannular air scoop 25 in line with the fuel tank 8. As can be seen fromFIGS. 1 and 2, the air is scooped between an external tubular structure1 and a two-way tapering ramp 9 whose geometry defines the features ofthe air scoop 25.

FIG. 2, which is a cross-sectional view of the missile in the plane ofthe scoop, shows how the ramp 9 is made. As shown in FIG. 2, this rampconsists of three segments 10 made for example of a composite casting.These three segments 10 of ramp 9 fulfill three functions: they definethe opening 11 of the air scoop 25, which is bounded by said segments 10and said external structure 1; secondly, they provide, together with theclearances 12 between the tank 8 and the ramp 9 and with the openings 13between said segments 10, a boundary layer trap; and thirdly, theycontribute to centering and securing the external structure 1 on thefuel tank 8, by means of the projections 14 on said segments 10 servingas bearings thereof.

The combustion chamber 2 is secured and centered within the externalstructure 1 at the rear of the missile by means of struts 15. Asindicated by FIGS. 3 and 4, the width of said struts 15 is not constant,but changes between the axial position represented in cross section inFIG. 3, in which plane are contained the first additional exhaustoutlets 6 provided in combustion chamber 2, and the axial positionfurther downstream represented in cross section in FIG. 4 containing themost downstream additional exhaust outlets 6. Said struts 15 have arelatively substantial width in the plane of FIG. 3. Said width iscalculated so that the available cross section 16 for passage of thesecondary air flow running between the combustion chamber 2 and theexternal structure 1, and thereby already preheated, provides a sonicflow. Between the radial planes illustrated in FIGS. 3 and 4 severalsets of additional exhaust outlets 6 serve, during the cruising stage,to inject exhaust gas into the secondary air stream formed betweenexternal structure 1 and combustion chamber 2. The width of said skids15 is reduced in the plane of FIG. 4 and is calculated to ensure thatthe cross section 16 available for the downstream runout of the exhaustgas is such as to also have a sonic mixture of combusted gas andsecondary air flow.

Downstream from the plane of FIG. 4, the exhaust gas/secondary air flowmixture expands in a sort of central body 18. Said central body 18 formsa basically conical fairing connecting the rear necking of thecombustion chamber 2 with the tapering out portion of the main axialnozzle 3 associated with the combustion chamber 2.

The furthest downstream part of the external structure 1 (FIG. 5) can beprovided with casings 19 of heat insulating material to house thecontrol surfaces servomotors in the case said control surfaces must belocated at the back of the missile.

The lower half-cross-section of FIG. 1 shows the state of the combustionchamber 2 prior to the acceleration stage, with the booster propellant 4masking the air inlets 5 as well as the additional exhaust outlets 6,said outlets being further sealed by means of plugs 7. The otherhalf-cross-section of FIG. 1 shows the combustion chamber 2 configuredfor the ramjet cruising mode, with the air inlets 5 and the additionalexhaust outlets 6 for injecting the combustion gases into the secondaryair stream now open.

During the acceleration stage, the air inlets 5 are obstructed by plugs7 bearing against the inside wall of the combustion chamber 2. Saidplugs are naturally ejected in the transition between the two propulsionmodes. Any suitable well known blow-out plug that will be released fromthe outlet 6 due to surrounding conditions may be used. The additionaloutlets 6 closing plugs 7 are ejected for example, by means of pressurewhen the solid propellant in the chamber 2 has been consumed. Inaddition, the ejection of plug 7 may be facilitated for example by meansof springs at the end off the propellant combustion tail.

According to one specific embodiment of the invention, the combustionchamber 2 is fabricated by filament winding, for example of phenolicsilica, and the fibers wind around inserts placed about the winding coreat the location of the air inlets 5. The additional outlets 6 can bemade in the same way. However, if the wall of the combustion chamber 2is reinforced in the area 17 where the additional outlets 6 are to beprovided, said outlets will be more easily provided by machining.

The operation of the missile according to the invention during ramjetcruising will be more readily understood with the help of FIG. 6. Thisfigure schematically illustrates with a lengthwise half-cross-sectionalview, the external structure 1, the fuel tank 8, the air inlets 5, theexhaust outlets 6, the main nozzle 3 and the body 18. It can be clearlyseen from the configuration in the figure that only part (αQa) of theair flow Qa drawn by air scoop 25 is introduced into the combustionchamber 2 via said air inlets 5, α having a value less than 1. Theremainder (1-α)Qa of the ram air Qa makes up the secondary air whichflows between the combustion chamber wall and the external structure 1and does not take part in combustion.

Similarly, part βQb(β<1) of the exhaust gas Qb exhausts axially via themain nozzle 3, while the remainder thereof (1-β)Qb exhausts via theadditional exhaust outlets 6 and mixes with the flow of secondary air(1-α)Qa. If the external structure 1 is extended downstream from thenozzle 3 far enough, it is possible to achieve mixing of all of theexhaust gas with the secondary air, the flow βQb of gas issuing from themain nozzle being thoroughly mixed with the flow {(1+α)Qa+(1-β)Qb} ofgas running out on the outside of said nozzle via an annular opening 26.

The dual-flow mode of operation according to the invention, ie.involving a secondary air flow (1-α)Qa running between the combustionchamber 2 and the external structure 1, which then mixes with at least apart (1-β)Qb of the exhaust gas, provides a number of advantages overconventional single-flow operation.

The primary advantage stems from the reduction of the velocity of thegases in chamber 2. This not only limits pressure drop and theconvection coefficient at the wall, but, more importantly, promotesthorough combustion due to a longer stay in the chamber, the physicallength of said combustion chamber 2, being defined in practice by thevolume of booster propellant, being in fact roughly the same as for asingle-flow ramjet design. Thus, good flame stabilization as well asgreater efficiency of combustion result from the low velocity in thechamber.

Another advantage is provided by the possibility of efficiently coolingthe shell of chamber 2, for the coefficient of external convection tothe non-combusted air is significantly greater than the coefficient ofinternal convection with the hot gases. Since the calories passingthrough the wall of the combustion chamber contribute to heating andaccelerating the secondary air flow, nothing prevents the walls of thechamber from being made thin and heat-conducting, providing they aremade from an oxidation-resistant material. This is advantageous to theextent that any reduction in thickness of the chamber walls enhances themissile's overall performance by freeing more space for boosterpropellant and/or ramjet fuel.

Moreover, in a dual-flow type ramjet engine, the combustion chambershell is under balanced pressure. The internal pressure differential inthe cruising mode becomes nil upstream and remains small downstream.

Furthermore, it deserves to be emphasized that during the accelerationstage with a conventional integral booster ramjet the combustion chamberair inlets are necessarily closed and consequently the missile's ram airaugments drag, by ΔF, for no useful purpose. The increase in drag on adual-flow ramjet during the acceleration stage on the other hand islimited to αΔF, α being a coefficient less than 1 standing for theproportion of air injected through the inlets 5 during the cruisingstage. Compared with a conventional ramjet engine therefore, thisinvention enables recovery of an amount of thrust (1-α)ΔF which is addedto the thrust of the solid propellant drive during the missile'sacceleration stage.

To summarize, a propulsion system or drive according to the presentinvention makes it possible, as with nozzle-less booster ramjets, not tojettison any heavy, compact body, while at the same time enabling aperformance comparable to that of ejectable nozzle ram rockets.

A second embodiment of the invention is depicted in FIGS. 7 and 8according to which opening of the air inlets 105 and of the additionalexhaust outlets 106 for combustion gas is done automatically thanks tothe sliding action imparted to the frame of the combustion chamber.

The drive system illustrated in FIGS. 7 and 8 comprises a nozzle 103specifically designed for acceleration stage operation and attached tothe missile's external structure 101, said latter structure withstandingboth internal pressure and flight stresses. A combustion chamber 102,initially containing the booster propellant, is positioned with respectto said external structure 101 by means of struts 109 establishing afree space between the external structure 101 and the wall of chamber102. Chamber 102 can slide backwards when during the booster combustiontail the resultant of the pressure forces on the head of the chamberbecomes greater than the resistance of a shearable tie. A tank 108located upstream from the chamber 102 contains the solidself-combustible fuel for the ramjet cruising stage. Said tank or store108 ignites from proximity alone and contributes to the impulsing of thebooster. In ramjet-type operation, the gases are fired spontaneously atcontact with the oxygen. No special ignitor is therefore required forthe cruising stage.

The booster operates in the same way as in the first embodimentdescribed hereinabove. An ignitor placed in said nozzle 103 startscombustion of the solid propellant placed in the combustion chamber 102.

As the missile accelerates, the ram air pressure increases and when thedirection of the resultant of the booster's internal pressure and of theexternal air pressure on the body of the combustion chamber 102reverses, during the combustion tail, said resultant eventually becominggreater than the strength of a shearable tie serving to initiallyimmobilize said combustion chamber 102, said justmentioned chamberslides back to open annular spaces 105 and 106 at the front and the backrespectively, so respectively allowing air entry to the chamber 102 andexhaust of the spent mixture. The transition between the two propulsionmodes (rocket to ramjet) thus differs from that of the first-mentionedembodiment. During the cruising stage, the very rich combustion productsgenerated in the chamber 102 are partly mixed with the secondary airflow in the free space 126 forming an open annular chamber between theexternal structure 101 and the main nozzle 103 of combustion chamber102, downstream from the annular space 106 forming an additional nozzle.The servomotors driving the control surfaces 122 can be housed in add-oncasings 120 of the external structure 101, in the space forming a mixingchamber.

The self-pyrolyzing solid propellant tank 108 communicates with thecombustion chamber 102 via an axially located outer opening 121 providedin the rear end of said tank 108. The nose cone 123 located ahead oftank 108 includes a point 124 which, together with the front end of theexternal structure 101, establishes an annular front air scoop 125. Thisconfiguration has the advantage of ensuring a good aerodynamicperformance and preserving the missile's symmetry of revolution. The airinlet 105 into the combustion chamber 102 which is formed during thecruising stage is annular. The cross-sectional area of air inlet 105 isroughly half that of the combustion chamber 102. The widening ofcross-sectional area accordingly formed has the effect of recycling airto stabilize combustion.

The main nozzle 103, the rear end fitting 110 of the combustion chamber102 which bears against the nozzle 103 during the acceleration stage andthe respective end fittings 111 and 112 of the front of the combustionchamber 102 and the rear of the tank 108 which cooperate to seal thecombustion chamber 102 during the acceleration stage, are fabricatedwith short fibers, such as for example, molded short silica fibers in aphenolic resin matrix, thus avoiding any problems or differentialexpansion as would be likely to occur with metal parts.

In the embodiment of FIGS. 7 and 8, the rear end fitting 110 of thecombustion chamber has a cylindrical section 115 of revolution about theaxis of the missile. In starting position, said section 115 surrounds amatching cylindrical section 116 of the nozzle 103 and applies to saidsection 116 in an airtight manner, a seal 113 being furthermore providedbetween the coaxial cylindrical sections 115 and 116 such that said twosections form a means 107 of sealing the additional annular outlet 106.In the transition from the acceleration stage to the cruising stage,said outside cylindrical section 115 simply slides over said insidecylindrical section 116, thus clearing an additional annular exhaustoutlet 106. At the front of combustion chamber 102, said end fitting 111is provided with a cylindrical flange of revolution 117, which instarting position engages in a mating slot 118 of the rear end fitting112 of tank 108, a seal 114 being arranged between the flange 117 of endfitting 111 and the slot 118 of end fitting 112. The slot 118 and flange117 have opposite faces parallel to the missile centerline so as toenable smooth sliding in the transition from the acceleration to thecruising mode.

For example, the ratio of the cross sectional area of the air scoop 125to that of the reference cross section consisting of the cross sectionalarea of the combustion chamber 102 can be in the neighborhood of 0.4,whereas the ratio of the cross section of the air inlet 105 to that ofsaid combustion chamber 102 is about 0.5. The ratio of the throat crosssection of main nozzle 103 to same said combustion chamber cross sectioncan be about 0.06 and the cross section of the additional exhaust outletor outlets 106 can be about twice that of said nozzle 103 throat,whereas the cross-sectional area for the flow of secondary air betweenthe external structure 101 and the chamber 102 can be on the order of1.4 times the cross-sectional area of the main nozzle 103.

What we claim is:
 1. An integral booster ramjet missile drive comprisinga single combustion chamber shared by a first, acceleration stage and asecond, ramjet cruising stage, said chamber having a casing with a rearnecked portion, said chamber housing a stored solid propellant used formissle acceleration, a converging-diverging nozzle optimally dimensionedfor acceleration stage propulsion and at least one air inlet designed toopen at the end of said acceleration stage to enable enough air to enterthe combustion chamber to at least compensate for a drag force on themissile by a high-speed ejection of gases resulting from the combustionof ram air with the stored solid propellant, said drive furthercomprising at least one additional exhaust outlet forming an additionalnozzle located at the downstream end of the combustion chamber, andmeans of closing said at least one additional exhaust outlet throughoutthe initial acceleration stage, such that said at least one additionaloutlet contributes, together with said converging-diverging nozzle, toexhausting the combustion gases during the ramjet cruising stage.
 2. Aramjet missile drive as in claim 1, wherein said additional exhaustoutlets are provided in the back wall of the combustion chamber toprovide a substantially axial additional exhaust.
 3. A ramjet missiledrive as in claim 1, wherein said additional exhaust outlets areprovided in an annular, rear lateral part of said combustion chamber. 4.A ramjet missile drive as in claim 1, further provided with an externalstructure surrounding said combustion chamber and extending forward overat least part of the length of the missile body so as to provide apermanent air scoop at the front of the missile and ensure ram air flowto the combustion chamber via said air inlets while maintaining thecasing of the combustion chamber under balanced pressure during theramjet cruising stage, such that during said cruising stage only a part(αQa) of the ram air (Qa) taken in by the missile's air scoops isintroduced into the combustion chamber via said air inlets while theremaining air flow (1-α)Qa is discharged behind the missile viapurposely provided outlets without having passed through the combustionchamber but having nevertheless been mixed with at least that portion(1-β)Qb of the exhaust gas exiting from the combustion chamber via saidadditional exhaust outlets.
 5. A ramjet missile drive as in claim 4,wherein said missile's external structure is extended backwards beyondthe main nozzle to enable mixing of the secondary air flow with all ofthe exhaust gas from the combustion chamber.
 6. A ramjet missile driveas in claim 4, further provided with struts between the casing of thecombustion chamber and said external structure, said struts being shapedso that the air flow cross section between said external structure andsaid chamber casing provides for sonic flow of both the air (1-α)Qaupstream from said additional exhaust outlets and the air and combustiongas mixture (1-β)Qb+(1-α)Qa downstream from said additional exhaustoutlets.
 7. A ramjet missile drive as in claim 6, the combustion chamberwhereof has a thin, heat-conducting wall.
 8. A ramjet missile drive asin claim 7, wherein said air scoops are not sealed off downstream duringthe acceleration stage to enable rearward discharge of the ram air.
 9. Aramjet missile drive as in claim 8, wherein a central body is provided,connecting said rear necked portion of the combustion chamber to anoutlet plane of the divergent portion of the converging-diverging nozzleconstituting a main axial nozzle, said central body forming a conicalfairing defining a divergent nozzle wherein the air and combustion gasmixture (1-α)Qa+(1-β)Qb is allowed to expand.
 10. A ramjet missile driveas in claim 9, wherein said air inlets as well as said additionalexhaust outlets are sealed off during the acceleration stage by specialplugs designed to be ejected at the end of said acceleration stage. 11.A ramjet missile drive as in claim 9, the combustion chamber whereof isdesigned to be automatically displaced by the pressure applied by thegases, axially backwards over a predetermined distance, at the end ofthe acceleration stage, thus opening said air inlets and/or additionalexhaust outlets provided in the combustion chamber.
 12. A ramjet missiledrive as in claim 11, the sliding combustion chamber whereof isinitially held immobilized by means of a shearable tie and thecombustion chamber whereof is so shaped as to ensure that at the end ofthe acceleration stage, the direction of the axial resultant of thepressure forces on said chamber reverses and said resultant becomesgreater than the strength of said shearable tie.
 13. A ramjet missiledrive as in claim 12, wherein said main nozzle and the end fittings ofthe combustion chamber and fuel tank are made by molding short fibers.